1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically for an air cooled large highly twisted and tapered turbine blade for an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A heavy duty large frame industrial gas turbine (IGT) engine is a very large engine with large turbine rotor blades. Current IGT engines include cooling for typically the first and second stage turbine vanes and blades. The later stage airfoils (vanes and blades) in the turbine do not require cooling because the hot gas stream temperature has dropped well below the melting temperatures of these airfoils. However, future IGT engines will have higher turbine inlet temperatures in which the third and even the fourth stage turbine rotor blades will require cooling in order to prevent significant creep damage. These hot turbine blades are under very high stress loads from rotating within the engine and therefore tend to creep of stretch from long period of operation. Creep issues are especially important for the lower sections of the blades because the lower section not only must provide structural support for the lower section of the blade but also for the upper section of the blade. Thus, internal cooling circuitry will be required in these blades.
Because of the increased spanwise length of these larger turbine rotor blades, the blade have a very high level of twist and taper for aerodynamic reasons. One prior art method of cooling a large turbine rotor blade is shown in U.S. Pat. No. 6,910,864 issued to Tomberg on Jun. 28, 2005 and entitled Turbine bucket airfoil cooling hole location, style and configuration. The cooling circuit for this blade includes drilling radial holes into the blade from the tip to the root. Limitations of drilling long radial holes from both ends of the airfoil section of the blade increases for a large highly twisted and tapered blade airfoil because the radial holes will not line up from the root to the tip. A reduction of the available cross sectional area for drilling radial holes is a function of the blade twist and taper. Higher airfoil twist and taper yield a lower available cross sectional area for drilling radial cooling holes. Cooling of the large, highly twisted and tapered blade by this process will not achieve the optimum blade cooling effectiveness required for future low flow cooling engines. It is also especially difficult to achieve effective cooling for the airfoil leading and trailing edges. Thus prevents higher turbine inlet temperatures for a large rotor blade cooling design that uses drilled radial cooling holes.